AU2003287542A1

AU2003287542A1 - Laminar flow wing for transonic cruise

Info

Publication number: AU2003287542A1
Application number: AU2003287542A
Authority: AU
Inventors: James D. Chase; Ilan Kroo; Richard R. Tracy
Original assignee: Aerion Corp
Current assignee: Aerion Corp
Priority date: 2002-11-07
Filing date: 2003-11-06
Publication date: 2004-06-03
Anticipated expiration: 2023-11-06
Also published as: CA2503270A1; US20040129837A1; EP1583690A2; AU2003287542B2; EP1583690A4; BR0316038A; ATE455034T1; EP1583690B1; WO2004043780A3; US7000870B2; WO2004043780A2; JP2006517162A; BR0316038B1; DE60331004D1; CA2503270C; JP4520306B2

Abstract

In combination, an aircraft wing and fuselage, comprising the wing having camber at or near the wing leading edge which has blunted sharpness and low sweep angle, and the fuselage having indentation along the wing side thereof, and lengthwise of the fuselage.

Description

WO 2004/043780 PCT/US2003/035393 LAMINAR FLOW WING FOR TRANSONIC CRUISE BACKGROUND OF THE INVENTION This invention relates generally to an aircraft wing adapted for s subsonic cruise conditions very near Mach One, and more particularly to a laminar wing adapted for such transonic cruise conditions. Modern subsonic aircraft designed for efficient cruise at high subsonic speed have moderate to highly swept wings (300 to 450) with moderately high aspect ratio (5 to 8), supercritical airfoils and t/c (thickness/chord) ratios as thin as possible, o10 typically about 10 to 12% for structural weight considerations. The latest fighters on the other hand, have much thinner airfoils, typically about 3 to 5% t/c as a compromise in order to also have supersonic dash capability, with similar leading edge sweep, but low aspect ratio (3 to 5), such that the trailing edge is nearly unswept. 15 Prior patent art of one of the present inventors (Tracy) describes a wing for efficient supersonic cruise which has a relatively unswept ( 300) and sharp "supersonic" leading edge, and t/c less than about 2% as a spanwise average, except in the immediate vicinity of the area-ruled fuselage. See U.S. Patents listed below. This wing achieves lower supersonic drag compared to the typical swept or delta 20 wing traditionally selected for efficient supersonic cruise by achieving extensive areas of laminar boundary layer coverage, which provide nearly an order of magnitude reduction in associated skin friction. The extremely low thickness is required to limit wave drag due to thickness to acceptable levels, and necessitates a low aspect ratio (2.5 to 3) for structural weight reasons. The sharp-edged airfoil is biconvex or a 25 minor modification thereof for low wave drag due to thickness and favorable pressure gradient. The low sweep is necessary to limit boundary layer crossflow to "subcritical levels" to prevent premature transition from laminar to turbulent boundary layer flow caused by so-called crossflow instability.

WO 2004/043780 PCT/US2003/035393 2 SUMMARY OF THE INVENTION It is, a major object of the invention to provide an improved aircraft wing adapted for aircraft transonic cruise conditions. The wing improves over the 5 wings described in Tracy U.S. Patents 6,149,101 and 5,897,076, and is useful for provision of improved transonic cruise efficiency, slightly below Mach 1. The wing is further characterized by more efficient cruise at higher subsonic speeds than can be achieved with the highly swept wings presently used for high subsonic cruise speeds. A further object is to provide an aircraft 10 wing and fuselage comprising a) said wing having camber (downward curvature) at or near the wing leading edge which has blunted sharpness, b) a fuselage having indentation along the wing side thereof, and lengthwise of the fuselage, 15 The fuselage and other components are typically "area-ruled" for a speed near Mach 1, (e.g. Mach 0.95), rather than for cruise at a supersonic Mach number as was done in the Tracy patent disclosing a wing for supersonic cruise aircraft. In addition the thickness-chord ratio is optimized for the transonic cruise condition. The thickness ratio is somewhat greater than the 2% average (along the 20 span) of the supersonic wing, but still much less than that of the swept wings used for high subsonic speed flight. The improved wing is further characterized by the following factors: Much thinner than current subsonic cruise wings, and at 2 to 4% t/c is similar to or thinner than current fighters. 25 Much lower leading edge sweep than current fighters, and at about 200 sweep or less is similar to, or less swept than the Tracy previously patented laminarsupersonic wing. Slightly blunted leading edge, and WO 2004/043780 PCT/US2003/035393 3 specially cambered and contoured airfoil (compared to the sharp, approximately biconvex supersonic airfoil of the prior Tracy patents). Low to moderate aspect ratio as a compromise between weight and cruise efficiency. 5 Locally thickened inner portion associated with "area ruled" fuselage. The airfoils are configured to create a "favorable" pressure gradient (pressure decreasing monotonically with distance aft from the leading edge) on the upper and 10 lower surfaces over the maximum fraction of chord at the lift coefficients associated with the design ciuise flight envelope, consistent with maximum practical critical Mach number at ihese conditions. Whereas the onset of compressibility generally creates local shock waves and unfavorable spanwise and chordwise pressure gradients causing transition on conventional subsonic laminar flow wings, the described wing 15 utilizes the compressibility effect at transonic Mach numbers to enhance the favorable pressure gradients needed for extensive laminar flow. A secondary consideration is achieving the highest practical maximum stall lift coefficient for the trailing edge flap system selected. Leading edge flaps, or devices are not necessary but desirable if the device doesn't introduce a disturbance capable of causing premature boundary layer 20 transition from laminarity to turbulence ("tripping" the boundary layer) when retracted at cruise conditions. The reasons for improved efficiency and other benefits associated with the thin unswept wing compared to the highly swept wings currently used for high subsonic cruise include: 25 * Reduced skin friction because of much more extensive laminar flow than possible with swept wings, as the result of low crossflow and extensive regions of favorable chordwise pressure gradient. * Onset of compressibility drag rise is WO 2004/043780 PCT/US2003/035393 4 "delayed" to higher subsonic Mach number with the very thin unswept wing, than for a typically thicker swept or delta wing. SThe unswept wing has improved takeoff and landing performance in the form of a lower s angle of attack than the swept wing, even if the latter has equal aspect ratio. The combined result of such effects is that the increase in total inviscid drag coefficient due to compressibility effects is delayed to higher Mach number very near Mach One, for the thin unswept wing. Even more important, the thin unswept o10 wing's extensive laminar flow near Mach one, which is not feasible for the swept wing, substantially reduces its skin friction drag, which is the key to its improved flight efficiency near Mach One compared to the delta or swept wings presently used for cruise at high subsonic speed. Further characteristics are: fuselage indentation along the wing side, 15 the wing at each side of the fuselage having generally trapezoidal configuration; the wing having camber at or near the wing leading edge, which has camber (droop) and blunted sharpness for transonic cruise; said wing having maximum thicknesses to chord length ratio (t/c) which varies with location from the wing root to the tip, the ratio of said maximum thickness to chord length at said locations varying from about 20 3.5% near the root to about 1.5% near the tip; wing thickness tapers forwardly and rearwardly from the zone of maximum thickness, with generally convex upper and lower surfaces, designed to maintain a smooth, favorable (decreasing) pressure gradient from the preferably slightly blunted leading edge, as far aft along the wing as practical at design flight conditions; wing root length subtending fuselage indentation; 25 wing leading edge sweep less than about 20 degrees, or as necessary to limit crossflow pressure gradients to levels which will not cause premature boundary layer flow transition from turbulent to laminar. Engines may be located in nacelles on either sides of the fuselage, or in association with the wing; or may be integrated into WO 2004/043780 PCT/US2003/035393 5 the fuselage. In ay case the effect of such engines is taken into account when the aircraft is "area ruled" as referenced above. Similar considerations apply to the tail surfaces of the aircraft, so as to achieve reduced drag and delay compressibility drag rise to higher subsonic speeds than 5 possible with typically thicker, swept tail surfaces. These and other objects and advantages of the invention, as well as the details of an illustrative embodiment, will be more fully understood from the following specification and drawings, in which: 10 DRAWING DESCRIPTION Fig.: 1 is a plan view of an aircraft incorporating the invention; Fig. 2 is a graph with abscissa correlated to the Fig. 1 aircraft length; and 15 Figs. 3-5 are airfoil sections taken on lines A-A, B-B and C-C, respectively of Fig; 1. Fig. 6 is a graph comparing the drag characteristics of an aircraft incorporating the invention compared to a typical swept wing aircraft. Fig. 7 is a graph illustrating the pressure distributions of an airfoil 20 section of the aircraft incorporating the invention at two Mach numbers. Fig. 8 magnifies the profile of the airfoil's slightly blunted leading edge. DETAILED DESCRIPTION 25 In the drawings, an aircraft 10 incorporating the invention has a fuselage 11, a wing 12 defined by left and right wing sections 12a and 12b; a tail at 13 WO 2004/043780 PCT/US2003/035393 6 having left and right sections 13a and 13b, and a vertical section 13c; and jet engines within nacelles 14 and 15. The latter are located near the inboard ends of wing trailing edges 12c and 12d and forward of the tail 13, as shown. The fuselage has width w, at location 34 proximate the inboard ends of the wing leading edges 12e and 5 12f; and width w, at a location 35 proximate such inboard ends of the trailing edges 12c and 12d, where: w, > w 2 , and the fuselage width between nacelle locations 14 and 15 decreases as shown. The fuselage and tail have locations 16 and 17 near the leading an trailing edges, o10 respectively of tail 13. Fig. 2 is a graph showing the aircraft cross sectional area, normal to the aircraft longitudinal axis. Note that between nacelle locations 14 and 15 (as referred to) the fuselage area decreases, with area contributions of the fuselage and wings as designated varying as shown at 18 and 19. Note also that between zones 33 15 and 22 (at or proximate the aft ends of the nacelles), the area contribution of the fuselage and nacelles vary as shown at 20 and 21. Note further that between zone 16 and 17 (as referred to) at or proximate the forward and aft ends of the tail, the area contribution of the fuselage and tail vary as shown at 24 and 25. 2o Referring to Figs. 1 and 3, the wing airfoil section 12g at station A - A (chordwise section) is shown to have leading edge blunted sharpness at 26, with camber, for high subsonic cruise condition. Figs. 4 and 5 show the airfoil section 12h and 12i at stations B-B and C-C (chordwise sections) with leading edge blunted sharpness 25 at 26' and 26", with camber, for high subsonic cruise. In Fig. 3, the maximum thickness to chord length ratio t/c is about 3.5%; that ratio in Fig. 4 is about 2.5%; and that ratio in Fig. 5 is about 1.5%.

WO 2004/043780 PCT/US2003/035393 7 Analysis shows that a wing configured according to the present invention will have a drag characteristic vs. Mach number much superior to a current swept wing in the high subsonic regime between Mach .9 and 1.0, as shown in Fig. 6. In Fig. 7, pressure distributions (in the form of pressure coefficients, 5 with negative values representing pressures below ambient and positive values pressures above ambient) are shown for upper and lower surfaces for an airfoil section shown at sections A-A of Figure 1. At Mach .8, below the design point of the wing, very sharp positive and negative pressure gradients are generated immediately aft of the leading edge. These sharp spikes in the pressure gradients would be 10 expected to cause transition from laminar flow, at least on the upper surface (upper curve). As speed increases to the Mach .95 design point, however, the compressibility effect is seen to modify the pressures into the more gradual negative gradients needed to foster extensive laminar flow on both upper and lower surface aft to about 70% of chord for the case shown. 15 Reduced drag between Mach 0.8 and 0.95 is the result of increased laminar flow, made possible by extensive favorable pressure gradients on both upper and lower surfaces as shown in Fig. 7 for cruise at Mach .95. The drag increase above 20 Mach 0.95 is caused by the increase in wave drag with Mach number overriding the drag reduction from the increasing extent of laminar flow. Fig. 8, illustrates a magnification of the leading edge. The relative "bluntness" of the leading edge is related by the ratio of the leading edge diameter "H" to airfoil maximum thickness "t". The "h/t" ratio is small, with ranges from 25 12 % to 3 % which in conjunction with the contours behind the leading edge maintains acceptable chordwise pressure gradients.

Claims

1. In combination, an aircraft wing and fuselage, comprising a) said wing having camber at or near the wing leading edge which has blunted sharpness, b) the wing having leading edge sweep less than about 200, c) the wing having thickness to chord ratio less than about 3 % as a spanwise average, and d) said fuselage having indentation along the wing side thereof, and lengthwise of the fuselage.

2. The combination of claim 1 wherein the wing at each side of the fuselage including said indentation has generally trapezoidal configuration.

3. The combination of claim 1 including a tail at the aft end of the fuselage, there being engine nacelles at opposite sides of the fuselage, and located between the wing and tail.

4. The combination of claim 1 wherein the wing has maximum thicknesses at locations along the wing from root to outer tip length, the ratio of said maximum thickness to chiord length at said locations varying from about 3.5 to about 1.5.

5. The combination of claim 4 wherein at each of said locations the wing tapers forwardly and rearwardly from the zone of maximum thickness, with generally convex upper and lower surfaces designed to maintain a smooth, favorable pressure gradient from the slightly blunted leading edge condition, to aft locations along the wing at subsonic design flight conditions. WO 2004/043780 PCT/US2003/035393 9

6. The combination of claim 1 wherein at each of said locations, the wing tapers forwardly and rearwardly from said zone of maximum thickness, throughout the wing length.

7. The combination of claim 1 wherein the wing at each side of the fuselage has a root length which subtends said indentation at said side.

8. The combination of claim 2 wherein the wing has leading edge sweep such as necessary to limit crossflow pressure gradients to levels which will not cause premature boundary layer transition at the design flight conditions.

9. The combination of claim 1 wherein the wing leading edge has blunted sharpness along substantially its entire length where the bluntness at each spanwise station is about 1/2% to 3 % of the maximum airfoil thickness at said station.

10. In combination, an aircraft wing and fuselage, comprising a) said wing having maximum thickness extending spanwise, said maximum thickness decreasing from a primary wing zone proximate the fuselage indentation to a secondary wing zone at a selected distance from the fuselage centerline, b) said fuselage having indentation along the wing side thereof, and lengthwise of the fuselage, c) said maximum thickness to chord radio, t/c remaining less than about 3.5% from said secondary zone to the wing tip.

11. The combination of claim 10 wherein said wing has camber at or near the wing leading edge which has blunted sharpness. WO 2004/043780 PCT/US2003/035393 20

12. The combination of claim 10 wherein the wing at each side of the fuselage including said indentation has trapezoidal configuration.

13. The combination of claim 11 including a tail at the aft end of the fuselage, there being engine nacelles at opposite sides of the fuselage, and located between the wing and tail.

14. The combination of claim 11 wherein said maximum thicknesses are associated with wing chord locations, at each of which the wing tapers forwardly and rearwardly, from the zone of maximum thickness with convex upper and lower surfaces designed to maintain a smooth, favorable pressure gradient from the slightly blunted leading edge, as far aft along the wings surface as feasible at subsonic design flight conditions.

15. The combination of claim 12 wherein the wing at each side of the fuselage has a root length which subtends said indentation at said side.

16. The combination of claim 13 wherein the tail tapers forwardly and rearwardly from zone of maximum thickness, with generally convex upper and lower surfaces designed to maintain a smooth, favorable pressure gradient condition from the slightly blunted leading edge, to locations as far aft along the tail as possible at subsonic design flight conditions.

17. In combination, an aircraft wing and fuselage, comprising a) said wing having camber at or near the wing leading edge which has blunted sharpness, b) the wing having leading edge low sweep angularity characterized in that crossflow instability is reduced to a subcritical level over the majority of the wing. WO 2004/043780 PCT/US2003/035393 11 c) the wing having thickness to chord ratio less than about 3 % as a spanwise average, and d) said fuselage having indentation along the wing side thereof, and lengthwise of the fuselage.